Load stability system

ABSTRACT

A heavy lift helicopter supporting a load suspended on a long cable is automatically controlled to minimize pendular motion of the load beneath the helicopter. The inner control loop of an automatic flight control system of a helicopter has an additional input which is the summation of a function of the rate of change of the angle between the cable and vertical, and the lag filtered rate of change of the angle, in both the lateral (roll) and longitudinal (pitch) directions of the aircraft. The sense of the rate of lag rate inputs is such as to cause the helicopter to move so as to counteract pendular motion of the load and support cable.

United States Patent [1 1 Fowler et al.

[ 1 LOAD STABILITY SYSTEM [451 Sept. 4, 1973 3,240,447 3/1966 Olshavsen244/77 D [75] ;E: g,g;: i West 3 2:? Primary ExaminerMilton Buchler Conroe o 0 Assistant Examiner-Stephen G. Kunin M PPT, t ,7, or V. vAttorneyMelvin Pearson Williams [73] Assignee: United AircraftCorporation, East HiEPFfSFFhEP m V [22] Filed: Aug. 5, 1971 57 ABSTRACT[2]] Appl. No.: 169,197 A heavy lift helicopter supporting a loadsuspended on a long cable is automatically controlled to minimizependular motion of the load beneath the helicopter. 'g The inner controlloop of an automatic flight control [58] Fie'ld 244;" 13 77 R system ofa helicopter has an additional input which is 44/77 6 137 the summationof a function of the rate of change of the angle between the cable andvertical, and the lag filtered rate of change of the angle, in both thelateral [5 6] References c'ted (roll) and longitudinal (pitch)directions of the aircraft. UNITED STATES PATENTS The sense of the rateof lag rate inputs is such as to 2,873,075 2/1959 Mooers et al 244/l7.l3cause the helicopter to move so as to counteract pen- 2 dular motion ofthe load and support cable. 3,055Z2l4 9/1962 McLane 244/77 D X 6 Claims,5 Drawing Figures JUMM/A/G AfZ/flfflfi 6 fifiidffi/t zf I i a @4 74 anyg /76// a p/m m 2%;

MRI 2 Bf 2 My wk PAIENTEDSEP 4191s I LOAD STABILITY SYSTEM BACKGROUND OFTHE INVENTION 1. Field of Invention This invention relates to aircraftcontrol systems, and more particularly to an automatic loadstabilization system for a cable supported load.

2. Description of the Prior Art The heavy lift helicopter is gainingacceptance as a means of moving relatively large and heavy objects whichare suspended therebeneath. Such a helicopter acts as a moveable cranein the sky. Typical examples of such use include the loading andunloading of ships, offshore oil operations, the building of skyscrapers, transmission line construction, assemblage of modularapartment buildings, and transporting miscellaneous material overuntravelable terrain. In military usage, loads are carried over enemyterritory, and ,over jungle, swamps, and other impassable terrain; also,armament may be moved very swiftly toward a battelfront therebyenhancing military logistics.

It is imperative that the helicopter not be rendered significantlyunstable while in flight since its ability to remain aloft might therebybe impaired. Therefore, the effect which the load has upon thehelicopter has to be predictable, rather constant as a function of time,and within the aerodynamic limits of the helicopter. Hopefully, the loadshould exert little more than a downward gravitational pull on theaircraft.

In accepted techniques for utilizing a heavy lift helicopter as a crane,a support cable extends from a winch, and the helicopter is caused tohover with substantially no motion or acceleration in any direction,while the cable is lowered and attached to the load, and the load isthereafter winched straight up to the helicopter. In this fashion, theload is carried very close to the helicopter so that the two exhibitmechanical properties of substantially a unitary body. However, thistype of maneuvering is very difficult (if not impossible) to achieveunder instrument flying conditions, where the visibility is very poor,or in high winds. As an example, consider the lifting of a heavy objectout of the hold of a ship. The object may be stabilized and guided fromthe ship until it clears the ship, provided the helicopter can bemaintained in a relatively stable position. However, maneuvering of thehelicopter so as to maintain its attitude naturally adjusts itsposition, and vice versa, which tends to affect the direction of pull bythe support cable on the load. This effect can be minimized by using arelatively long cable. On the other hand, once the load clears the ship,it is then free to swing until such time as it has been winched intocontact (or nearly so) with the helicopter. But if the cable is long,the load may begin to swing, with a pendular motion, even while it is onthe way up and while it might still come into contact with thesuper-structure of the ship. On the other hand, if a short cable isused, the removal of the load from within the ship is very difficulteven though it may quickly be winched to the helicopter once it is clearof the hold.

Depending upon weather conditions and the flight maneuvering which isrequired, it may be dangerous to have a load winched directly to thebottom of a helicopter. That is, the dynamic motion of the helicoptermay result in the load literally colliding with the helicopter.

Therefore, more recently developed techniques utilize a very long cablewith the load suspended therefrom. In some crane type aircraft known tothe art, the helicopter is controlled by a second pilot sitting in anaft seat facing the winch where he can observe the load; in allaircraft, the pilot maneuvers the aircraft so as to stabilize the load.Under manual control, loadsupporting helicopters are limited in speed,and suffer reduced maneuverability.

SUMMARY OF INVENTION The object of the present invention is to provideload stabilization for a load suspended below an aircraft on a loadsupport cable.

According to the present invention, pendular motion of load suspended ona cable below an aircraft is corrected by causing the aircraft to movein a direction related to functions of the rate of change of angle ofthe support cable with respect to the vertical. In accordance with theinvention in one form, the functions of the angle rate approximates, ona short term basis, the angle minus the rate of change of the angle.

In further accord with the present invention, the afore-mentionedstabilization is achieved utilizing the summation of the rate of changeof cable angle and the lag of the rate of change of cable angle tomodify th flight control parameters of the aircraft.

In one embodiment of the invention, a load supported from a helicopteris stabilized by modifying the inner control loop of the automaticflight control system of the helicopter by adding thereto the rate ofchange of cable angle and the lag, or washed out integral, of the rateof change of the angle of the cable with respect to vertical. I

The present invention provides correct anti-pendular corrective motionsto an aircraft. The invention may be embodied utilizing technologyreadily available in existing flight control systems, and is easilyappended to existing helicopter automatic flight control systems. Theinvention is relatively inexpensive to provide, and renders theutilization of heavy lift helicopters practicable in a wide variety ofoperations, such as crane type operations, under extended conditions offlight.

The foregoing and other objects, features and advantages of the presentinvention will become more apparent in the light of the followingdetailed description of a preferred embodiment thereof as illustrated inthe ac companying drawing.

BRIEF DESCRIPTION OF THE DRAWING FIG. 1 is a simplified front elevationview of a helicopter supporting a load swinging in pendular motion tostarboard;

FIG. 2 is a simplified front elevation view of a helicopter rolled toport to overcome or correct the situation of FIG. 1;

FIG. 3 is a simplified side elevation view of a helicopter in forwardflight illustrating a lag of the load behind the helicopter in steadystate forward flight condition;

FIG. 4 is a simplified, illustrative schematic block diagram of thelongitudinal (or pitch) cyclic pitch control channel in accordance withthe present invention; and

FIG. 5 is a simplified chart illustrating time histories of theLaPlacian functions utilized in the embodiment of FIG. 4.

DESCRIPTION OF THE PREFERRED EMBODIMENT Referring now to FIG. 1 ahelicopter supports a load 22 by means of a support cable 24 suspendedbelow the helicopter. As seen in FIG. 1, the helicopter 20 issubstantially horizontal and may be assumed to be in straight lineflight (toward the viewer) or hovering. However, the load 22 isexhibiting pendular motion, and at the moment is moving to starboard ofthe helicopter 20. In accordance with one aspect of the presentinvention, it has been found that the proper way to correct for pendularmotion is to first fly the helicopter 20 to starboard so as to attemptto fly in the same direction as the motion of the load, and then, as isillustrated in FIG. 2, to tilt the helicopter 20 in a direction to causeit to move away from the load. As described in more detail with respectto FIG. 3 hereinafter, this motion, generally opposite to the simpleharmonic motion of the load, is adjusted by utilizing the rate of changeof the angle between the support cable 24 and the vertical, and the lagof that rate, in such a fashion as to cause corrective motion of theaircraft to damp out pendular motion of the load with respect thereto.As an example, consider the case where the aircraft is hovering andthere is a gust of wind which blows the load 22 to the starboard asshown in FIG. 1. Subject to the inertia of the load, the rate of changeof the angle of the cable 24 with respect to the vertical becomes highinitially, then, in accordance with the physics of simple harmonicmotion, the load tends to reach its maximum position at which the rateis zero. Initially, when the rate is very high, this is utilized as asignal of polarity to cause the helicopter to initially move tostarboard to tend to stay over the load; but as the rate decreases, andthe angle increases, eventually they cancel each other and then theangle becomes more of a factor as the rate of change of the angleapproaches zero. The net polarity then reverses so as to tilt theaircraft as illustrated in FIG. 2 to tend to pull the load in adirection opposite to the manner in which the load is swinging in simpleharmonic motion; in other words as seen in FIG. 2 the motion of theaircraft tends to damp the kinetic energy of the load. By using a lagfilter, however, in the steady state (after an interval equal to severaltime constants of the lag filter involved) it is possible for the loadto hang at a steady non-zero angle with respect to the vertical, asillustrated in FIG. 3, with no steady state control signal. This isnecessary since, in straight line flight, with or without acceleration,the load will lag behind the aircraft as illustrated in FIG. 3, due toaerodynamic drag forces. A similar condition appertains while executinga turn since it is necessary to overcome aerodynamic drag, centrifugalforce necessary to accelerate in the turn, and lag of the load behindthe aircraft. All of these functions are accommodated in one embodimentof the invention as illustrated in FIG. 4.

In FIG. 4, the aircraft itself has an inertial body 20 and is shownseparate from the various blocks of the longitudinal (or pitch) cyclicpitch control channel for illustrative purposes. As is known, the pilotoperates a combined lateral (roll) and longitudinal cyclic pitch controlstick 30 which is mechanically coupled (32) to a mechanical summingactuator 34 which controls the cyclic pitch of the blades 36 of the mainrotor of the helicopter 20. As is known, motion of the stick 30 from theposition shown in solid lines to the position 30 shown in dotted lineswould cause the aircraft to pitch in an upwardly direction, climb, anddecelerate in forward flight. The actual mechanical motion at the output38 of the summing actuator 34 is modified from the mechanical motion ofthe input 32 by a mechanical input to the summing actuator 34 from anelectromechanical actuator 40. The electromechanical actuator 40 derivesan input from a summing network 42 which in turn is responsive to a pairof amplifiers 44, 46 having gains of K and K,, respectively. Theamplifiers 44, 46 are in turn responsive to a pitch gyro 48 and pitchrate gyro 50, respectively. The apparatus 40-50 just described comprisesthe normal automatic flight control system (AFCS) of the aircraft, whichis sometimes referred to as the inner control loop of the aircraft sincethe functions thereof do not affect the motion of the stick itself, butmerely modify the results of stick motion.

In FIG. 4, a sign convention is defined that an upward pitch as shown istaken as a positive pitch angle and it is assumed that, as a result ofsuch positive pitch, the output of the pitch gyro 48 and the output ofthe pitch rate gyro 50 will both be of such a polarity as applied to thesumming network 42 and electromechanic actuator 40 to counteract thepositive pitch of the aircraft. Thus, these are shown as negative inputsto the summing network 42.

In accordance with the present invention, the control system of theaircraft is modified so as to stabilize the load 22 in such a fashion asto tend to maintain the cable 24 vertical (though not necessarilyperpendicular to the frame of reference of the aircraft). According tothe invention, a cable angle rate gyro 52 is mounted on the aircraft insuch a fashion that it measures only the rate of change (de/dt) of theangle e of the cable with respect to the vertical in a fore-aft plane,and not with respect to the attitude of the aircraft, nor side-to-side.The output of the cable angle rate gyro 52 is applied to an amplifier 54with a gain of K the output of which is applied to the summing network42. The output of the cable angle rate gyro 52 is also applied to a lagcircuit 56 having a function K,/( trl-l where K, is the amplificationfactor, t is the time constant, and s is the La- Placian operator(equivalent to the time derivative). The time constant, t, is chosensuch that the function of the circuit 56 on a short term basis isequivalent to the angle e. It is to be noted that the polarity of theamplifier 54 as applied to the summing network 52 is negative (the sameas the gyros 48, 50) so that, as the load begins to swing and the rateof change of the angle 2 with respect to time (e) is high, the aircraftinitially receives an input of the opposite sense (causing it to pitchdown) so the aircraft would fly forward in the convention of FIG. 4 andthus attempt to stay over the load momentarily. However, after a shortperiod of time, the output of amplifier 56 begins to build up, as theoutput of the amplifier 54 decreases. Since the amplifier 56 is appliedto the summing network in the positive sense, it tends to overcome theoutput of the amplifier 54, and eventually causes a positive pitch ofthe aircraft so that the aircraft would tend to decelerate, therebyapplying an opposite, backward acceleration on the load 22 through thecable 24, to tend to dampen the pendular motion of the load. Toillustrate this combined action, FIG. 5 shows just a portion of a swingof the load from vertical, forwardly, to the point as shown in FIG. 4.This example is for a response of the load to a short wind gust, tendingto swing the load forward. The operation of the two amplifiers 54, 56 inresponse to the cable angle rate gyro is illustrated roughly therein,although not to any particular scale. As seen in FIG. 5, illustration Ashows the case where the cable angle builds up from zero (vertical) to avalue plus e (as shown in FIG. 4) at time T In this case the timeconstant, t, is taken to be relatively long with respect to a short termoscillation (short term in this case being equivalent to T IllustrationB shows that the rate of change of the angle, e, builds up from zero tosome maximum at about T and then decreases to zero at time T On theother hand, illustration C shows that the lag function, e/(ts-i-l is, inthe short term, essentially equal to the angle e (dependent upon theamplification factor I(,) so that it maximizes somewhere near time T andonce the rate of change of angle goes to zero (at time T it then decaysback toward zero. The combination of the two (which is the effect thatthe outputs of both amplifiers 54, 56 have at the summing network 42) isshown in illustration D. This is the control input which is applied tothe aircraft to overcome the effect of the pendular action of the load.Initially, it supplies a negative input which tends to cause theaircraft to fly over the load, then applies no input at time T and thenapplies a positive input which causes the aircraft to decelerate (asseen in FIG. 4), thereby acting against the swinging of the load so thatduring the period of time between T, and T it is causing the aircraft topull against the swinging of the load thereby dissipating the kineticenergy or dampening the harmonic motion of the load. It continues topull against the load tending to pull the load back under the aircraftfrom time T onwardly, in the simple case illustrated herein. Of course,in the real situation, the waveforms of FIG. 5 become complicated sincethe rate of change of angle will go negative after time T, as the loadcommences to swing back aft. This is further complicated by the factthat the aircraft itself is in motion and has not only the inputstending to stabilize the load but other inputs of the AFCS, and perhapsfrom the pilot. It is for this reason that only the simple portion ofthe cycle has been shown in FIG. 5.

The lateral (or roll) cyclic pitch control channel for a helicopter inaccordance with an embodiment of the present invention is identical tothe longitudinal (or pitch) control channel illustrated in FIG. 4, withthe exception of the fact that the constants are chosen differentlysince the aircraft response is different in pitch and in roll, and thetime constant might be slightly less since the aircraft can respond inroll more rapidly than it can in pitch. A similar sign convention isachieved simply by defining motion of the stick to starboard to bepositive; tilting of the aircraft downward on the starboard side to bepositive; and a cable angle which is to port of vertical to be positive.In the lateral controls, a cable angle sensor responsive only to rate ofchange of angle of the cable 24 from vertical in a port-starboard planeprovides the necessary rate signal.

As is known, the time constants of the amplifiers, the amplificationfactors and the polarity of inputs are chosen so as to suit the dynamicsof any given implementation of the present invention. For instance, astabilization analysis of a given aircraft with a typical de- 6 56. Asan example, a very short cable may require that polarity of the outputof either amplifier 54, 56 be reversed in the sign convention of FIG. 4.It is important to note that the choice of parameters including sign canbe determined from the standard stability analysis which is well knownto the art.

The embodiment of the invention described herein is related to a heavylift helicopter. However, it should be obvious that other aircraft maybe employed in a system incorporating the present invention. Althoughboth pitch and pitch rate gyros are shown, the invention may bepracticed utilizing only a pitch gyro and taking the derivativetherefrom for the gyro rate function. Similarly, if desired, cable anglemay be sensed with position sensors (such as potentiometers) and therate thereof provided by differentiating the output of the positionsensor. The invention may be embodied in other aircraft control systems:for instance, a more complete aircraft control system may includelateral and longitudinal accelerometers to assist in positioning theaircraft in a stable fashion, or may include a ground speed input tofacilitate accurate maneuvering of the aircraft.

Thus, although the invention has been shown and described with respectto a preferred embodiment thereof, it should be understood by thoseskilled in the art that the foregoing and various other changes andomissions in the form and detail thereof may be made therein withoutdeparting from the spirit and the scope of the invention.

Having thus described a typical embodiment of our invention, that whichwe claim as new and desire to secure by Letters Patent of the UnitedStates is:

1. Aircraft control system apparatus for stabilizing, against pendularmotion, a load suspended beneath the aircraft on a cable exhibitingpendular motion having a basic period including first intervalssubstantially characterized by relatively high velocity and relativelysmall angle with respect to the vertical, and including second intervalssubstantially characterized by rela-- tively low velocity and relativelylarge angle with respect to the vertical, comprising:

means for sensing motion of said cable with respect to vertical and forgenerating a manifestation as a function of said motion; and meansincluding first means responsive to said manifestation to cause saidaircraft to move in a'direction the same as the motion of the loadduring a substantial portion of said first intervals and includingsecond means also responsive to said manifestation to cause saidaircraft to move in a direction opposite to the motion of the loadduring a substantial portion of said second intervals. 2. Apparatusaccording to claim 1 wherein said manifestation generating meanscomprises:

means responsive to changes in the position of said cable with respectto vertical to generate a rate signal indication of the rate-of changeof the angle of said cable with respect to vertical; and meansresponsive to said rate signal to generate said manifestation as thedifference between the lag of said rate and said rate. 3. A system forstablizing a load suspended beneath an aircraft having a plurality ofcontrol surfaces for controlling the attitude and motion thereof, andincluding a flight control system for providing at least partial controlover said control surfaces, comprising:

a load support cable adapted to extend downwardly beneath said aircraftand to support a load thereon; means responsive to changes in theposition of said cable with respect to vertical to generate a ratesignal indication of the rate of change of the angle of said cable withrespect to vertical; and

cause motion of said aircraft so as to counteract pendular motion ofsaid load therebeneath. 4. A system according to claim 3 wherein saidadditional function is proportional to the lag of said rate. 5. Thesystem according to claim 3 wherein said additional function isproportional to the washed out integral of said rate.

6. The system according to claim 3 wherein said addimeans responsive tosaid rate signal indication to 10 tional function is provided by a lagfilter having a transvide a signal manifestation input to said flightcontrol system which is dependent upon said rate and an additionalfunction of said rate in a manner to fer characteristic proportional tol/(ts-l-l) where s is the LaPlacian operator and t is the time constant.

I. I I t

1. Aircraft control system apparatus for stabilizing, against pendularmotion, a load suspended beneath the aircraft on a cable exhibitingpendular motion having a basic period including first intervalssubstantially characterized by relatively high velocity and relativelysmall angle with respect to the vertical, and including second intervalssubstantially characterized by relatively low velocity and relativelylarge angle with respect to the vertical, comprising: means for sensingmotion of said cable with respect to vertical and for generating amanifestation as a function of said motion; and means including firstmeans responsive to said manifestation to cause said aircraft to move ina direction the same as the motion of the load during a substantialportion of said first intervals and including second means alsoresponsive to said manifestation to cause said aircraft to move in adirection opposite to the motion of the load during a substantialportion of said second intervals.
 2. Apparatus according to claim 1wherein said manifestation generating means comprises: means responsiveto changes in the position of said cable with respect to vertical togenerate a rate signal indication of the rate of change of the angle ofsaid cable with respect to vertical; and means responsive to said ratesignal to generate said manifestation as the difference between the lagof said rate and said rate.
 3. A system for stablizing a load suspendedbeneath an aircraft having a plurality of control surfaces forcontrolling the attitude and motion thereof, and including a flightcontrol system for providing at least partial control over said controlsurfaces, comprising: a load support cable adapted to extend downwardlybeneath said aircraft and to support a load thereon; means responsive tochanges in the position of said cable with respect to vertical togenerate a rate signal indication of the rate of change of the angle ofsaid cable with respect to vertical; and means responsive to said ratesignal indication to provide a signal manifestation input to said flightcontrol system which is dependent upon said rate and an additionalfunction of said rate in a manner to cause motion of said aircraft so asto counteract pendular motion of said load therebeneath.
 4. A systemaccording to claim 3 wherein said additional function is proportional tothe lag of said rate.
 5. The system according to claim 3 wherein saidadditional function is proportional to the washed out integral of saidrate.
 6. The system according to claim 3 wherein said additionalfunction is provided by a lag filter having a transfer characteristicproportional to 1/(ts+1) where s is the LaPlacian operator and t is thetime constant.